Typical inertial guidance systems employ various sensors to measure changes in motion. For example, a linear accelerometer is often used to measure acceleration along one linear axis. A combination of three separate linear accelerometers can measure acceleration in all three dimensions. These measurements of acceleration can then be integrated over time to determine the velocity and position of an object. Based on position and velocity readings, the object's trajectory is adjusted to arrive at a desired target. Inertial guidance systems place accelerometers on a platform also containing gyroscopes to measure or control the orientation of the accelerometers. The gyroscopes maintain the accelerometers in either a real physical or virtual computed inertial frame of reference by providing measures of the angular rotation in inertial space. The angular rotation information is then used to either control the orientation of the sensors so that they experience no rotation with respect to inertial space or to compute the orientation with respect to inertial space or to both control and compute the orientation. The type of inertial guidance system which controls the physical orientation of the sensors to remain rotationally fixed is commonly known as an inertially stabilized system. The type of inertial guidance system which computes the orientation of the sensors with respect to inertial space is commonly known as a strap-down system whether or not the unit remains rotationally fixed to the missile or aircraft. The control forces necessary to control the orientation of the gyroscopes and accelerometers can be obtained by providing appropriate torques to a gimbal set which contains the sensor cluster or to a freely suspended sphere which contains the instrument cluster. Applications of these inertial guidance systems can be found in many avionics systems such as missiles and commercial or military aircraft.
High accuracy inertial guidance systems such as those used in avionics systems can undergo high accelerations or input rates as well as experience gradual changes in offset or bias errors which are independent of the magnitude of the input. High accelerations or input rates lead to sensor errors that are often difficult to calibrate prior to flight and bias errors can trend after calibration to new values at flight. For example, errors in acceleration measurement can be due to accelerometer bias or high levels of acceleration driving scale factor nonlinearities. As the acceleration measurements are integrated to determine position and velocity, these acceleration errors will cause position and velocity errors that grow in time. Virtually all inertial guidance systems suffer from this problem of integration error growth.
For the reasons stated above and for other reasons stated below which will become apparent to those skilled in the art upon reading and understanding the present specification, there is a need in the art for a measurement unit which can reduce the errors in measurement due to both fixed and slowly changing errors.